This disclosure relates generally to drilling systems and methods for preparing composite material and substrate materials for assembly and, in particular, relates to tools and methods for deburring after drilling operations.
Composite components are being utilized in a wide variety of articles of manufacture due to their high strength and light weight. This is particularly true in the field of aircraft manufacturing. Typical materials used in the manufacture of composite components include glass or graphite fibers that are embedded in resins, such as phenolic, epoxy, and bismaleimide resins. A composite lamination can be built up by laying successive plies of fiber tows (e.g., carbon fiber tows preimpregnated with a thermoset epoxy resin) around a mandrel and then curing. As more advanced materials and a wider variety of material forms have become available, aerospace usage of composites has increased.
For example, composites are used in conjunction with metal substrates to form an assembly that may be used to construct a larger structure, for example, of an aircraft or other vehicle. The assembly may include a composite material and a structural metal substrate arranged in a stacked or layered orientation (referred to herein as a “stackup”). The composite material may, for example, be carbon fiber-reinforced plastic (CFRP) or other fiber-reinforced material. The structural metal substrate may, for example, be titanium, aluminum or steel. The metal substrate may be used to build a skeleton or frame, with the composite material attached to and covering the frame. For this reason the composite material is sometimes referred to as a skin. The metal and composite materials may be shaped, contoured, or curved into virtually any shape desired.
Of course the composite material and metal substrate have different physical attributes and properties, and exhibit different behavior in use. Due to those facts, attaching the composite material to the metal substrate can be challenging. For example, the materials may be joined to each another with a fastener that requires holes to be drilled in each respective material. Separate handling of the composite material from the metal substrate is undesirable. Especially for relatively large structures having many fasteners distributed over the structure, such as in the fabrication of an aircraft, avoiding separate drilling of the holes in each of the composite material and the metal substrate may result in appreciable reductions in production times and reduction in costs of fabricating the aircraft.
To avoid separate drilling, many machining applications involve drilling and/or reaming a hybrid stack-up of composite and metal materials. For example, certain aircraft require that a wing made from a composite material, such as CFRP, be joined to a titanium section of an aircraft body with fasteners that pass through holes made through the mating sections. When using fasteners to attach composite skins to metal substrates, coaxial holes must be drilled in both the skin and an underlying metal substrate. High-quality holes must be produced in such materials with dimensions within narrow tolerances.
The wing-to-body join task typically requires a three-step conventional drilling process comprising a pilot drill, followed by a step drill, followed by a finish diameter reamer. Frequently the reaming process is followed by a deburring operation. Various special tools are known for removing burrs from the circumferential edges surrounding openings of drilled holes and for adding chamfers thereto. In particular, mechanical hole-deburring tools are known which remove burrs on the front, back, or both sides of drilled holes in one pass, working from one side only.
The design of airframe structure dictates the elements in the stack. Metal components are often times “sandwiched” between CFRP components. The stack orientations are driven by structural loading requirements. The high-load areas at the wing-to-body interface typically have external metal components whereas the body section joins are mostly CFRP with interior metal components. Location and access are the primary drivers for determining from which direction one can approach the interface for deburring operations.
A known deburring blade has been used to perform metal material removal as part of a deburring operation within a drilled and/or reamed hole in a mixed material interface, e.g. CFRP/Ti or Ti/CFRP. The existing blade design provides for cutting force reaction in only a single direction. With the existing blade design, when attempting to deburr using the forward portion of a double-acting cutter tip, the cutting reaction forces “push” the blade back in the hole and a reduced amount of material is removed.
There is a need for a blade design that will enable loading of the cutter blade from either side of the cutter tip without any relative movement away from the metal interface as a result of cutting forces.